Vapor retention device

ABSTRACT

Embodiments of the present invention generally relate to a vapor retention device and methods of using a vapor retention device to manage propellant for upper stage space vehicles. The use of a vapor retention device, in combination with controlled acceleration, drives liquid propellant from a propellant supply line communicating with an upper stage main engine back into a propellant tank and establishes an insulating liquid/gas propellant interface that prevents the exchange of gaseous propellant across the interface.

FIELD OF THE INVENTION

Embodiments of the present invention generally relate to a vaporretention device associated with a cryogenic liquid tank and enginefeedline.

BACKGROUND OF THE INVENTION

There is an increasing demand for longer duration space transport atlower cost. Long duration missions are not possible with existingcryogenic launch vehicles. With expectations set on pushing humanitythroughout cislunar space and beyond, upper stage performance, includingpropellant management, must be enhanced.

Upper stage vehicles may utilize a single propellant or two propellants.An existing upper stage, such as Centaur®, utilizes liquid hydrogen andliquid oxygen at cryogenic temperatures. The propellants are storedseparately in lightweight stainless-steel tanks whose structuralrigidity is provided primarily by the pressure of the propellants withineach tank. Typically, the tanks are located coaxially, with the liquidoxygen tank positioned between the liquid hydrogen tank and the engines.This configuration requires the feedline for the liquid hydrogen tank torun around or through the liquid oxygen tank. As a cryogenic propellantwarms, including liquid hydrogen and oxygen, it changes phase to a gaswhich accumulates within the propellant tank. For example, while on theground, the cryogenic propellant within the hydrogen tank boils due toexternal heating and pressure within the tank increases. Pressure reliefvalves are used to vent the building pressure and liquid hydrogenpropellant is continuously added to the tank to offset the loss from thephase change. However, losses in space due to external heating cannot becounterbalanced by adding new liquid propellant. Sources of heatinginclude solar heating, heat transfer from warmer propellants (like LO2)to colder propellants (like LH2), as well as soak back heating fromengine burns during the course of a mission. Solar heating can warm theliquid propellant making the propellant too warm to burn or cause theliquid propellant to boil off into gas which is then unavailable toburn. One method of addressing solar heating is to reposition the upperstage such that the propellant tanks are shielded from direct solarexposure by other portions of the upper stage vehicle. In addition, anyengine burn generates heat within the engine and the surroundingstructures, including the propellant feedlines between the engine andthe fuel tanks. Even after engine shutdown, liquid propellant remainingin feedlines is subject to warming which, in turn, changes the liquidpropellant to gas. Furthermore, when the stage is under acceleration,buoyancy causes the gas to migrate through the feedline to thepropellant tank, further warming and converting liquid propellant to gasand draws cold propellant towards the bottom of the tank where theseheating sources cause more boiloff. Accordingly, fuel is lost andmission capabilities are degraded.

Devices exist to separate gas propellant from liquid propellant in afuel tank, however, these devices exist to ensure solely liquidpropellant is fed to the engine and not to prevent or reduceintermingling of liquid and gas propellant during engine operation.These devices maintain the liquid propellant in contact with the engineinlet at all times through surface tension and capillary action and donot prevent parasitic heating from continuing to convert liquidpropellant to gas and buoyancy from bringing fresh propellant to theheating source. For example, U.S. Pat. No. 5,901,557 entitled “PassiveLow Gravity Cryogenic Storage Vessel,” discloses a cryogenic propellantstorage vessel with a screen trap that divides the interior of thevessel into a liquid phase compartment and gas phase compartment. Theoverall purpose is to prevent the latter from entering the engine. Overthe course of a flight, ullage is formed and comingles with propellantused to fuel rocket burns. However, surface tension formed by liquidpropellant interaction with a vane assembly prevent the ullage or gasphase propellant from entering the engine. A complex heat exchangesystem is further positioned in the base and sump of the propellant tankto cool the propellant.

In U.S. Pat. No. 8,381,938 entitled “Propellant Tank for CryogenicLiquids,” the propellant tank includes an exterior mounted reservoir.The reservoir is in communication with the interior of the tank to allowliquid to flow into the reservoir from where it is then supplied to anengine. Gas is driven out of the reservoir and back into the tank. Thissystem allows migration of warmer gas through liquid propellant therebywarming the liquid propellant and converting some liquid propellant to agas phase.

U.S. Pat. No. 5,293,895 entitled “Liquid Management Apparatus forSpacecraft” discloses a liquid propellant tank that includes a fill,drain, and feedline extending from one end of the tank and a screenlocated proximate the feedline. The purpose is to enhance engineperformance at zero gravity. Capillary action is used to move propellantto the engine inlet. A screen is located near the engine inlet and iswetted with a liquid film as liquid propellant is withdrawn from thetank for purposes of supplying the engine. The purpose of the screen isto prevent gas existing within the main body of the propellant tank frompassing through the screen, entering the feedline and comingling withthe supply of propellant to the engine. Engine performance degradesappreciably if gas is comingled with liquid propellant.

SUMMARY OF THE INVENTION

According to embodiments of the present disclosure, methods andapparatus are provided to preserve liquid cryogenic propellant inconnection with the operation of a spacecraft.

According to aspects of the present disclosure, an upper stage spacevehicle is provided which utilizes a liquid cryogenic propellant systemand, more specifically, uses the formation of gaseous cryogenicpropellant to preserve liquid cryogenic propellant and thereby extendthe duration of a mission. A vapor retention device is used to capturegaseous propellant in the propellant supply line extending between aliquid cryogenic propellant tank and the main engine for the upperstage. The gaseous propellant is created from stationary liquidpropellant remaining in the supply line following an engine burn. Thesupply line and other surrounding structures, are heated during anengine burn. When the engine burn is halted and the flow of liquidpropellant ceases, the heated supply line warms the stationary liquidpropellant causing it to transition to a gas phase of the propellant.The formation of the gaseous propellant is used to drive the remainingliquid propellant in the supply line into the propellant tank where itis preserved for future use. Simultaneously, settling motors are used tosettle the upper stage, including the liquid propellant within thepropellant tank. The settling motors maintain the upper stage under asmall acceleration which positions the liquid propellant against oneside of the vapor retention device. Allowing the liquid propellant topass through the vapor retention device but preventing the gaseouspropellant from passing through. The vapor retention device uses liquidsurface tension effects to form an interface along the surface of thevapor retention device to prevent gaseous propellant from passing intothe propellant tank and warming the stored liquid propellant.Acceleration levels are selected for the coast period which allowgaseous propellant bubbles to stay attached to the vapor retentiondevice and maintain the liquid/gas interface.

According to aspects of the present disclosure, a cryogenic propellantsystem for an upper stage space vehicle is disclosed. In at least oneembodiment, the upper stage space vehicle has an engine, a firstpropellant tank containing a first cryogenic propellant and having apropellant port for filling and withdrawing the first cryogenicpropellant. It also has a second propellant tank containing a secondcryogenic propellant, the second propellant tank is typically coaxiallyaligned with the first propellant tank although the tanks may bepositioned in any other arrangement. A first propellant supply line isin fluid communication between the propellant port of the firstpropellant tank and the engine associated with the upper stage spacevehicle and a first length of the first supply line extends through oraround the second propellant tank and terminates at the main engine. Avapor retention device is positioned in the first propellant tankproximate the propellant port and restricts the flow of gas from thefirst propellant supply line into the first propellant tank during lowacceleration periods of flight and permits the flow of liquid propellantfrom the first propellant tank to the engine through the firstpropellant supply line during high acceleration periods of flight.

According to aspects of the present disclosure, a method for preservinga cryogenic propellant in connection with the operation of a spacevehicle is provided. The method includes providing a space vehiclehaving at least one engine, at least one propellant tank containing acryogenic liquid propellant, a propellant port associated with thepropellant tank, a propellant supply line having a first end connectedto the propellant port and a second end in communication with the atleast one engine, and a vapor retention device positioned proximate thepropellant port either in the propellant tank, in the supply line or inthe port itself. The method further includes, upon shutting down the atleast one engine, settling the propellant in the propellant tank andapplying a small acceleration to the upper stage while simultaneouslyusing the transition of liquid cryogenic propellant to gaseous cryogenicpropellant to force liquid cryogenic propellant from the supply lineinto the propellant tank. It should be appreciated that this method maybe practiced in association with each of multiple liquid cryogenicpropellant tanks. The method further includes using the vapor retentiondevice to prevent the gas phase propellant from entering the propellanttank.

According to aspects of the present disclosure, the vapor retentiondevice is a plate with a plurality of apertures. The plate may have aflat or curved surface or both. It may be dome shaped, box shaped,conical or some other shape. Factors that influence the size of theapertures include propellant settling acceleration levels selected foruse during a coast periods, propellant supply line geometry to allow forgood flow during high g operation (engine burns) and slug flow duringlow g operation (coasts).

According to aspects of the present disclosure, acceleration of theupper stage can be created in many ways. The method described hereinwill consist of pulsed settling where the time averaged acceleration iswhat is important. However, controlled constant thrust or centrifugalrotational acceleration methods create the same effect. Pulsed settlingcan be time averaged to create the effect of constant accelerationthrough the selection of the pulse duration which is known to those ofskill in the art. Therefore, through selection of the pulsed thrustperiod and knowledge of the stage mass, acceleration is controlled.Consequently, mission simulations are required to predict the stage massas a function of time or measure acceleration precisely in order tosequence the correct settling levels for a particular mission.

The term “upper stage” as used herein means any stage or spacecraftcarried into space by a launch vehicle. A launch vehicle may carry morethan one upper stage into space.

The phrases “at least one”, “one or more”, and “and/or”, as used herein,are open-ended expressions that are both conjunctive and disjunctive inoperation. For example, each of the expressions “at least one of A, Band C”, “at least one of A, B, or C”, “one or more of A, B, and C”, “oneor more of A, B, or C” and “A, B, and/or C” means A alone, B alone, Calone, A and B together, A and C together, B and C together, or A, B andC together.

Unless otherwise indicated, all numbers expressing quantities,dimensions, conditions, and so forth used in the specification andclaims are to be understood as being modified in all instances by theterm “about”.

The term “a” or “an” entity, as used herein, refers to one or more ofthat entity. As such, the terms “a” (or “an”), “one or more” and “atleast one” can be used interchangeably herein.

The use of “including,” “comprising,” or “having” and variations thereofherein is meant to encompass the items listed thereafter and equivalentsthereof as well as additional items. Accordingly, the terms “including,”“comprising,” or “having” and variations thereof can be usedinterchangeably herein.

It shall be understood that the term “means” as used herein shall begiven its broadest possible interpretation in accordance with 35 U.S.C.§ 112(f). Accordingly, a claim incorporating the term “means” shallcover all structures, materials, or acts set forth herein, and all ofthe equivalents thereof. Further, the structures, materials, or acts andthe equivalents thereof shall include all those described in the summaryof the invention, brief description of the drawings, detaileddescription, abstract, and claims themselves.

These and other advantages will be apparent from the disclosure of theinvention(s) contained herein. The above-described embodiments,objectives, and configurations are neither complete nor exhaustive. TheSummary of the Invention is neither intended nor should it be construedas being representative of the full extent and scope of the presentinvention. Moreover, references made herein to “the present invention”or aspects thereof should be understood to mean certain embodiments ofthe present invention and should not necessarily be construed aslimiting all embodiments to a particular description. The presentinvention is set forth in various levels of detail in the Summary of theInvention as well as in the attached drawings and the DetailedDescription and no limitation as to the scope of the present inventionis intended by either the inclusion or non-inclusion of elements,components, etc. in this Summary of the Invention. Additional aspects ofthe present invention will become more readily apparent from theDetailed Description, particularly when taken together with thedrawings.

The above-described benefits, embodiments, and/or characterizations arenot necessarily complete or exhaustive, and in particular, as to thepatentable subject matter disclosed herein. Other benefits, embodiments,and/or characterizations of the present disclosure are possibleutilizing, alone or in combination, as set forth above and/or describedin the accompanying figures and/or in the description herein below.However, the Detailed Description, the drawing figures, and theexemplary claims set forth herein, taken in conjunction with thisSummary of the Invention, define the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

Those of skill in the art will recognize that the following descriptionis merely illustrative of the principles of the invention, which may beapplied in various ways to provide many different alternativeembodiments. This description is made for illustrating the generalprinciples of the teachings of this invention and is not meant to limitthe inventive concepts disclosed herein.

The accompanying drawings, which are incorporated in and constitute apart of the specification, illustrate embodiments of the invention and,together with the general description of the invention given above andthe detailed description of the drawings given below, serve to explainthe principles of the invention.

FIG. 1 is a flow chart representing a cycling of a propellant systemassociated with a launch vehicle.

FIG. 2 is a section view of a first embodiment of a dual propellantsystem for a launch vehicle, with one liquid propellant being added to apropellant tank, for example, prior to launch.

FIG. 3A is a section view of the embodiment of FIG. 2 during engineburn, where propellant is supplied to the engine(s).

FIG. 3B is a section view of the embodiment of FIG. 3A, post launch,where the engines are shut down and the launch vehicle is entering acoast mode.

FIG. 3C is a section view of the embodiment of FIG. 3B, still in a coastmode but later in time compared to FIG. 3B.

FIG. 3D is a section view of the embodiment of FIG. 3C, still in a coastmode but later in time compared to FIG. 3C.

FIG. 3E is a section view of the embodiment of FIG. 3D, still in a coastmode but later in time compared to FIG. 3D.

FIG. 3F is a section view of the embodiment of FIG. 3E, still in a coastmode but later in time compared to FIG. 3E.

FIG. 4A is a perspective view of one embodiment of a vapor retentiondevice according to the present disclosure.

FIG. 4B is an enlarged cross-section view of an aperture of the vaporretention device of the embodiment of FIG. 4A.

FIG. 5 is a section view of the embodiment of FIG. 3F, with thepropellant system preparing for an engine burn.

FIG. 6 is a section view of the embodiment of FIG. 5, with the engineactively burning propellant.

The drawings are not necessarily to scale, and various dimensions may bealtered. In certain instances, details that are not necessary for anunderstanding of the invention or that render other details difficult toperceive may have been omitted. It should be understood, of course, thatthe invention is not necessarily limited to the particular embodimentsillustrated herein.

DETAILED DESCRIPTION

Although the following text sets forth a detailed description ofembodiments according to the present disclosure, it should be understoodthat the legal scope of the description is defined by the words of theclaims set forth at the end of this disclosure. The detailed descriptionis to be construed as exemplary only and does not describe everypossible embodiment since describing every possible embodiment would beimpractical, if not impossible. Numerous alternative embodiments couldbe implemented, using either current technology or technology developedafter the filing date of this patent, which would still fall within thescope of the claims.

The orientation and directions as used herein are relative to thedrawings as illustrated. Therefore, it should be appreciated that theterms “above,” “below,” “top,” “bottom,” “horizontal,” or “vertical” areused to describe the relative location of different parts of the novelmechanism following launch, the position of the spacecraft may no longerremain vertical but may have other orientations. Thus, the novelmechanism may be oriented differently in flight, but the relativeposition of the novel mechanism is as described.

FIG. 1 is a flow chart representing typically recurring states of anupper stage spacecraft over the duration of a mission. Initially, alaunch vehicle and associated upper stage are on a launch pad preparingfor ascent. Liquid propellant is continuously fed into propellant tanks,and the tanks are continuously vented to equalize pressure within thepropellant tanks due to boil off. Prior to lift off, connections to thelaunch vehicle and upper stage are disconnected, liquid propellant isdelivered to the launch vehicle engines from one or more propellanttanks and the engines ignite and burn the propellant. The launch vehicleascends through the atmosphere. Acceleration from the main engines ofthe launch vehicle is typically between 0.5 and 4.0 g. At apredetermined point, the launch vehicle separates from the upper stage.The launch vehicle returns to Earth. Following separation, the mainengines on the upper stage ignite and, at 10, the upper stage continuesits ascent. Liquid propellant from one or more tanks on the upper stagesupplies the main engines. At 20, the engines of the upper stage areshut down and a period of coasting begins. Settling motors are used tosettle the movement or sloshing of the propellant within each tankcaused by vibrations generated by the engine burn of both the launchvehicle and upper stage engines. At 30, the liquid propellant in thetanks of the upper stage has settled and the upper stage is in a longcoast phase. During the coast period, the engines cool. Following engineshut down, residual propellant in the supply lines between thepropellant tanks and the engine is likely to turn to gas due to heatfrom the engine(s) and external sources warming the propellant supplyline to a temperature that cause liquid propellant to boil off. Theupper stage will also be repositioned to counteract external solarheating. Settling motors are used to accomplish any repositioning and tore-settle propellant due to such repositioning. At 40, the upper stageprepares to end the period of coasting. At 50, the engines are preparedfor a burn. Any gas phase propellant within the supply line and engineis evacuated and the supply lines are filled with liquid propellant. At10, the upper stage main engines are again ignited and propellant isburned. This cycle may be repeated as necessary to position the upperstage for mission objectives.

FIGS. 2-5 and 7 are cross section views of one embodiment of apropellant system 200 designed for a launch vehicle or upper stagespacecraft. A two-propellant system is illustrated, however it should beunderstood that the present disclosure applies to single or multiplepropellant systems. In these illustrations, the upper tank 210 storesliquid hydrogen and the lower tank 212 stores liquid oxygen. Differentpropellants may be used. A central vent line 214 extends from a valve216 through the lower liquid oxygen tank 212 and into the upper liquidhydrogen tank 210. A liquid hydrogen supply line 218 extends from thehydrogen tank 210 to the main engines (not shown). The vent and supplyline may also go around the lower tank. Valves 220 are positionedbetween the main engines and the hydrogen tank 210 to control the flowof propellant to the engines. A portion of the hydrogen supply line 218traverses the liquid oxygen tank 212 and surrounds the vent line 214forming an annular space 222 for the flow of cryogenic liquid hydrogento the engines. A vapor retention device 224 is positioned proximate theinterface between the supply line 218 and the hydrogen tank 210. Itshould be understood that in the case of a single propellant vehicle thesupply line would not extend through another tank but would otherwiseserve the same purpose.

Turning to FIG. 2, a propellant system 200 for a launch vehicle or upperstage is shown as it may exist prior to launch. The upper tank 210includes liquid hydrogen 230 and the lower tank 212 included liquidoxygen 232. Gaseous hydrogen or ullage 228 is forming in the upperportion of the upper tank 210 due to boil off. The gas is vented out ofthe tank through valve 216 to maintain tank pressure and proper levels.As also illustrated, valves 220 are open and liquid hydrogen is beingsupplied to the upper tank 210 through the hydrogen supply to replaceliquid propellant losses due to boil off Although not shown, the sameactions are occurring with respect to the liquid oxygen tank. The liquidoxygen 232 within tank 212 also generates ullage due to boiloff of theliquid oxygen. Although not illustrated, the oxygen tank 212 wouldsimilarly include a liquid oxygen supply line, a gaseous oxygen accessline and associated valves to control the flow of liquid and gaseousoxygen into and out of the tank 212.

FIG. 3A corresponds to state 10 in FIG. 1. Specifically, liquid hydrogen230 is being supplied to the main engines (not shown) for engine burn.The flow of the liquid hydrogen is represented by the arrows in supplyline 218. Liquid oxygen 232 is also being supplied to the engines, butthe supply line is not shown. During engine burn, the heat produced bythe engine heats the surrounding structures, including supply line 218and valves 220. However, the liquid propellants are moving from thetanks 210 and 212 at a sufficient rate to prevent boil off.

FIG. 3B illustrates the propellant system 200 following engine shutdown. This corresponds to state 20 in FIG. 1. At this point, the liquidpropellant is no longer flowing in the supply line 218 but is nowstagnant. Valves 220 are closed. However, the liquid hydrogen 230 andliquid oxygen 232 are sloshing within the propellant tanks 210 and 212,respectively, due to the vibration generated by the operation and shutdown of the main engines. Heating of the propellant sump and supplylines (indicated by the arrows) initiates boil off of the liquidpropellant at the hot surfaces. Low settling levels and heating in thesupply line 218 closest the engine cause gaseous hydrogen 234 initiallyto form along the inner wall of the supply line 218 closest to the hotengine. Gaseous bubbles form, coalesce and grow larger. Surface tensioneffects dominate. The annular space 222 is still filled with liquidhydrogen 230 at this point in time.

As illustrated in FIG. 3C, settling of the cryogenic propellants(hydrogen and oxygen) has occurred due to operation of the settlingmotors. The upper stage is in a coast period. The gaseous hydrogenbubbles 234 continue to grow in the supply line 218 as more liquidpropellant boils off. Ultimately, slug flow occurs and the enlarginghydrogen gas bubble pushes liquid hydrogen 230 remaining in the supplyline back into the liquid hydrogen tank 210. This is sequentiallyillustrated in FIGS. 3B-3F. The slug flow of the gaseous hydrogen 234has a positive benefit. It acts similar to a piston, forcing liquidhydrogen 230 back into the liquid hydrogen tank 210 where the returnedliquid hydrogen is preserved in a liquid state for future use. Bychanneling the effects of boil off, liquid propellant may be preservedfor future use instead of transitioning to a gas phase that is notburnable by the main engines.

A vapor retention device 224 is positioned at the propellant supply port240 of the hydrogen tank 210. One embodiment of a vapor retention device224 is illustrated in FIGS. 4A and B. Another embodiment of a vaporretention device for alternate tank configuration would place the deviceat the tank exit independent of the orientation of the supply line. Thevapor retention device comprises a screen or membrane with openings orapertures sized to create a stable liquid-gas interface at theacceleration provided by the upper stage during a coast. Typically,settling thrusters or motors (not shown) provide the neededacceleration. This interface prevents the gaseous hydrogen 234 formed inthe supply line 218 from migrating into the liquid hydrogen tank 210 andprevents liquid hydrogen 230 from migrating into the supply line 218where heating sources would continue to create gaseous hydrogen 234 fromthe fresh liquid hydrogen 230. In addition, the hydrogen gas 234 trappedin the supply line 218 acts as an insulating barrier between the coldliquid propellants and the heating sources shown as arrows. Thus,instead of losing additional liquid propellant due to boil off,embodiments of the present disclosure harness the formation of gaseouspropellant in supply lines to push liquid propellant back intopropellant tanks and restrict the movement of the liquid propellantpreventing it from coming into contact with sources of heat which wouldcreate excessive boiloff. Settling acceleration of the upper stage isutilized to create the buoyant driving force which enhances the pistoneffect described above.

FIG. 3F illustrates a propellant system state where the supply line 218is empty of liquid propellant. This state corresponds to a coast state,such as at 30 in FIG. 1. Motion or movement of the upper stage iscontrolled by settling motors to maintain settling forces below athreshold value that allows the effects of surface tension within thegaseous hydrogen 234 at the interface with the vapor retention device224 to dominate. According to aspects of the present disclosure,maintaining an acceleration of between about 10⁻³ and 10⁻⁵ G's willallow the surface tension to dominate. This effect is illustrated by thearrows which represent a balanced or controlled level of settling thatmaintains the needed stable liquid-gas interface. As a result, thegaseous hydrogen is prevented from migrating through the vapor retentiondevice and entering the liquid propellant tank 210 and liquid isprevented from entering the supply line 218 where it would encounterdeleterious heating.

Those of skill in the art will know the physical force balance term“Bond Number” which can be utilized to determine the stability of aliquid gas interface. This parameter will indicate the maximum dimensionwhere a stable interface will exist at a certain acceleration.Consequently, by matching vehicle acceleration and critical dimensionsof the apertures in the vapor retention device a stable interface can becontrolled and utilized in the method shown herein. Because theacceleration and critical dimensions are inversely related, once theacceleration goes below that threshold the stable interface will remain,including down to 0 G conditions, unless dislodged by pressuredifferentials which occur, for example, when the upper stage mainengines operate and acceleration exceeds the threshold value. Heattransfer between the liquid oxygen tank 212 and the liquid hydrogen tank210 by way of the feedline 218 is reduced by removal of the liquidhydrogen from the annular space 222 since heat transfer is driven by atemperature difference and surface area. Liquid hydrogen, typicallymaintained at −420 degrees Fahrenheit, is held away from the liquidoxygen, typically maintained at −320 degrees Fahrenheit and other heatsources, such as the propellant sump and supply lines. The only contactarea between the cold liquid hydrogen and warm gas is across the stableliquid-gas interface in the VRD represented by arrows. The heat transferacross the liquid gas interface and along thin pieces of structure areorders of magnitudes lower than direct conduction from the heat sourceacross tank structure to liquid cryogens.

FIG. 4A illustrates an example of how a vapor retention device could bedesigned to create the propellant saving effects described herein.Apertures 240 are sized so that the critical dimension matches thechosen stage acceleration. That is at the chosen stage accelerationduring long coasts the liquid gas interface will remain stably attachedto holes of this size or smaller. According to aspects for the presentdisclosure, the holes preferably are on the order of 0.5-3.0 inches indiameter depending on the stage acceleration chosen for the long coast.Openings within this diameter range will work with a coast accelerationof between about 10⁻³ to 10⁻⁵ Gs. The shape of the hole and hole edgesalso play a role in promoting the liquid gas interface attaching to itwithout unduly hindering the flow once the vehicle enters state 50 and10. FIG. 4B illustrates one preferred example of a hole shape. Asillustrated, the vapor retention device has a first surface 242 and anopposed surface 244. With reference to FIG. 4A, the first surface 242 isthe inner surface and the second surface 244 is the outer or uppersurface. At the first surface 242, the edge 246 of the hole is sharpangle while at the outer surface 244 there is no discrete edge, butrather a gradual rounded transition 248 from the surface 244 to theinside surface 250 of the aperture 240. Liquid propellant would bepositioned on the side proximate first surface 242 and gaseouspropellant would be positioned on the side proximate second surface 244.The angle between the inside surface 250 and first surface 242 isapproximately 90 degrees but could be between approximately 90 and 45degrees, preferably approximately 70 to 90 degrees, and most preferablyapproximately 80 to 90 degrees. Up to a 5-degree change in the anglewould fall within these ranges. The sharp edge 246 enhances the abilityof the liquid to attach to the vapor retention device 240 and inhibitthe transfer of gas into a propellant tank. Alternative designs mayinclude variations of hole placement and vapor retention device shape toselect the angle between the centerline of the hole and the axialacceleration vector of the vehicle. For example, instead of a dome shapeas illustrated in FIG. 4A, the vapor retention device 224 could byconical, frusto-conical, pyramid or box shaped. Further considerationsto the vapor retention device include preventing large pressure drops atthe high flowrates to the engines during state 10. The quantity of holesrelative to the opening size of the holes must be sufficient to supplypropellant to the main engines as needed during prescribed burns.

FIG. 5 corresponds to state 40 in FIG. 1. Here, the upper stage orspacecraft is ending a coasting period. Settling of the liquidpropellant increases in the aft direction. This increases the pressureof the liquid propellant against the vapor retention device 224.Increased settling changes the balance of the surface tension forceswith other forces on the interface with the vapor retention device 224.Arrows represent the breakdown of the stable liquid-gas interface andthe start of liquid flow into the gas filled supply line 218. This flowwill start to chill the supply line in preparation for an engine burn.

FIG. 6 corresponds to state 50 of FIG. 1. Here, the upper stage ispreparing for an engine ignition and burn. Engine valves (not shown) andsupply line valves 220 are opened allowing the pressure in the liquidhydrogen tank 210 to push out the gaseous propellant 234 from the supplylines 218 and push in liquid propellant 230 to fill the supply linesincluding line 218 and finishes chilling the supply lines back to liquidhydrogen temperatures. Once gaseous hydrogen 234 is evacuated from theengine and supply lines and liquid hydrogen and oxygen fill the supplylines, the main engines are ignited. This state is illustrated in FIG.3A and corresponds to state 10 in FIG. 1.

While various embodiments of the present invention have been describedin detail, it is apparent that modifications and alterations of thoseembodiments will occur to those skilled in the art. However, it is to beexpressly understood that such modifications and alterations are withinthe scope and spirit of the present invention, as set forth in thefollowing claims. Further, the invention(s) described herein is capableof other embodiments and of being practiced or of being carried out invarious ways. It is to be understood that the phraseology andterminology used herein is for the purpose of description and should notbe regarded as limiting.

What is claimed is:
 1. A method for preserving a cryogenic propellant inconnection with the operation of an upper stage vehicle, comprising:providing an upper stage vehicle having at least one main engine, afirst propellant tank containing a first cryogenic liquid propellant, afirst propellant port associated with the first propellant tank, and afirst propellant supply line having a first end connected to the firstpropellant port and a second end in communication with the at least oneengine; covering the first propellant port with a first vapor retentiondevice positioned in either the first propellant tank or the firstpropellant supply line; supplying the first cryogenic liquid propellantto the at least one engine using the first supply line; consuming thefirst cryogenic liquid propellant by operating the at least one engine;following ceasing operation of the at least one engine, changing thephase of a first portion of the first cryogenic liquid propellant to agaseous phase in the first supply line; forcing a second portion of thefirst liquid cryogenic propellant remaining in the first supply linemove into the first propellant tank; forming an insulation barrier onthe engine side of the first vapor retention device, the insulationbarrier comprising the gaseous phase of the first portion of the firstcryogenic liquid propellant; maintaining an interface at the first vaporretention device between the first liquid cryogenic propellant and thegaseous phase of the first cryogenic propellant to prevent the gaseousphase of the first portion of the first cryogenic propellant within thefirst propellant supply line from entering the first propellant tank. 2.The method of claim 1, wherein maintaining the interface comprisesmaintaining a surface tension of the first liquid cryogenic propellanton the vapor retention device.
 3. The method of claim 2, whereinmaintaining the interface comprises controlling the acceleration of theupper stage vehicle.
 4. The method of claim 3, further comprisingaccelerating the upper stage at a rate of less than approximately 10-3 gto maintain the first propellant liquid interface on the vapor retentiondevice.
 5. The method of claim 4, wherein the acceleration of the upperstage does not exceed approximately 10-5 g.
 6. The method of claim 1,wherein changing the phase of the first cryogenic liquid propellant to agaseous phase propellant comprises heating the first propellant supplyline by burning first cryogenic liquid propellant in the at least oneengine.
 7. The method of claim 1, wherein forming an insulation barriercomprises maintaining the gaseous phase of the first propellant in thefirst propellant supply line.
 8. The method of claim 1, furthercomprising: providing a second propellant tank containing a secondcryogenic propellant, a second propellant port associated with thesecond propellant tank, and a second propellant supply line incommunication between the second propellant port and the at least oneengine; covering the second propellant port with a second vaporretention device positioned in either the second propellant tank or thesecond propellant supply line; supplying the second cryogenic propellantto the at least one engine using the second supply line; consuming thesecond cryogenic propellant during the operation of the at least oneengine; following the operation of the at least one engine, changing thephase of a first portion of the second cryogenic liquid propellant to agaseous phase in the second supply line; forcing a second portion of thesecond liquid cryogenic propellant remaining in the second supply linemove into the second propellant tank; forming an insulation barrier onthe engine side of the second vapor retention device, the insulationbarrier comprising the gaseous phase of the first portion of the secondcryogenic liquid propellant; maintaining an interface at the secondvapor retention device between the second liquid cryogenic propellantand the gas phase of the first portion of the second gaseous cryogenicpropellant to prevent the gas phase of the first portion of the secondcryogenic propellant within the second propellant supply line fromentering the second propellant tank.
 9. The method of claim 1, whereinthe vapor retention device comprises a body with a plurality ofsubstantially round holes having a diameter between 0.5 and 3.0 inches.10. The method of claim 9, wherein the body has a thickness, a firstsurface and a second surface, the plurality of holes extend between thefirst and second surfaces, and each hole has a side wall extendingbetween the first and second surfaces, and wherein the transitionbetween the first surface and the side wall of each hole is an angle andthe transition between the second surface and the side wall is rounded,and wherein the angle is approximately 90 degrees.
 11. A method forpreserving a cryogenic propellant in connection with the operation of anupper stage vehicle, comprising: providing an upper stage vehicle havingat least one main engine, a first propellant tank containing a firstcryogenic liquid propellant, a first propellant port associated with thefirst propellant tank, and a first propellant supply line having a firstend connected to the first propellant port and a second end incommunication with the at least one engine; covering the firstpropellant port with a vapor retention device positioned in either thefirst propellant tank or the first propellant supply line, the vaporretention device comprising a body having a thickness, a first surfaceand a second surface, a plurality of holes extending between the firstand second surfaces and having a diameter between 0.5 and 3.0 inches,each hole having a side wall extending between the first and secondsurfaces, and wherein the transition between the first surface and theside wall of each hole is an angle of approximately 90 degrees and thetransition between the second surface and the side wall is rounded;supplying the first cryogenic liquid propellant to the at least oneengine using the first supply line; consuming the first cryogenic liquidpropellant by operating the at least one engine; following ceasingoperation of the at least one engine, changing the phase of a firstportion of the first cryogenic liquid propellant to a gaseous phase inthe first supply line; forcing a second portion of the first liquidcryogenic propellant remaining in the first supply line move into thefirst propellant tank; forming an insulation barrier on the engine sideof the vapor retention device, the insulation barrier comprising thegaseous phase of the first portion of the first cryogenic liquidpropellant; maintaining an interface at the vapor retention devicebetween the first liquid cryogenic propellant and the gaseous phase ofthe first portion of the first cryogenic propellant to prevent the gasphase of the first portion of the first cryogenic propellant within thefirst propellant supply line from entering the first propellant tank.12. The method of claim 11, wherein maintaining the interface comprisesmaintaining a surface tension of the first liquid cryogenic propellanton the vapor retention device.
 13. The method of claim 12, whereinmaintaining the interface comprises controlling the acceleration of theupper stage vehicle.
 14. The method of claim 13, further comprisingaccelerating the upper stage at a rate of less than approximately 10-3 gto maintain the first propellant liquid interface on the vapor retentiondevice.
 15. The method of claim 14, wherein the acceleration of theupper stage does not exceed approximately 10-5 g.